Annular combustor with tangential cooling air injection

ABSTRACT

The combustion dynamics and efficiency of gas turbine having an annular combustor 26 provided with fuel injection nozzles 50 that inject fuel generally tangentially is improved by providing the walls 32, 34, 39 of the combustion 26 with cooling air film injectors 70, 86; 72, 88; 74, 90 at substantially equally angularly spaced locations about each such wall and which are oriented to generally tangentially inject a film-like air stream on the associated wall 32, 34, 39.

.Iadd.This application is a a reissue of Ser. No. 07/138,342 filed Dec.28, 1987, now U.S. Pat. No. 4,928,479. .Iaddend.

FIELD OF THE INVENTION

This invention relates to gas turbines, and more particularly, to animproved combustor for use in gas turbines.

BACKGROUND OF THE INVENTION

It has long been known that achieving uniform circumferential turbineinlet temperature distribution in gas turbines is highly desirable.Uniform distribution minimizes hot spots and cold spots to maximizeefficiency of operation as well as prolongs the life of those parts ofthe turbine exposed to hot gasses.

To achieve uniform turbine inlet temperature distribution in gasturbines having annular combustors, one has had to provide a largenumber of fuel injectors to assure that the fuel is uniformlydistributed in the combustion air. Fuel injectors are quite expensivewith the consequence that the use of a large number of them is noteconomically satisfactory. Moreover, as the number of fuel injectorsincreases in a system, with unchanged fuel consumption, the flow areafor fuel in each injector becomes smaller. As the fuel flow passagesbecome progressively smaller, the injectors are more prone to cloggingdue to very small contaminants in the fuel.

This in turn creates the very problem sought to be done away withthrough the use of a number of fuel injectors. In particular, a fouledfuel injector will result in a non uniform turbine inlet temperature inan annular combustor with the result that hot and cold spots occur.

To avoid this difficulty, the prior art has suggested that by and largeaxial injection using a plurality of injectors be modified to the extentthat such injectors inject the fuel into the annular combustion chamberwith some sort of tangential component. The resulting swirl of fuel andcombustion supporting gas provides a much more uniform mix of fuel withthe air to provide a more uniform burn and thus achieve morecircumferential uniformity in the turbine inlet temperature. However,this solution deals only with minimizing the presence of hot and/or coldspots when one or more injectors plug and does not deal with thedesirability of eliminating a number of fuel injectors to reduce costand/or avoiding the use of injectors having very small fuel flowpassages which are prone to clogging.

The present invention is directed to overcoming one or more of the aboveproblem.

SUMMARY OF THE INVENTION

It is the principal object of the invention to provide a new andimproved annular combustor for a gas turbine. More specifically, it isan object of the invention to provide such a combustor wherein thenumber of fuel injectors may be minimized and yet uniformcircumferential turbine inlet temperature distribution retained alongwith a minimization the possibility of the fuel injectors plugging.

An exemplary embodiment of the invention achieves the foregoing objectsin a gas turbine including a rotor having compressor blades ana turbineblades. An inlet is located adjacent one side of the compressor bladesand a diffuser is located adjacent the other side of the compressorblades. A nozzle is disposed adjacent the turbine blades for directinghot gasses at the turbine blades to cause rotation of the rotor and anannular combustor having spaced radially inner and outer, axiallyextending wails connected by a radially extending wall is disposed aboutthe rotor and has an outlet connected to the nozzle and a primarycombustion annulus remote from the outlet. A plurality of fuel injectorsto the primary combustion annulus are provided and are substantiallyequally angular spaced about the same. They are configured to injectfuel into the primary combustion annulus in a nominally tangentialdirection. Cooling air for one or more of the walls of the annularcombustor is introduced tangentially in a film-like fashion along theinterior side or sides of one or more of the combustor walls. The use ofa tangentially flowing film of cooling air serves to reduce the tendencyof injected fuel from moving in the axial direction allowing completeevaporation within the primary combustion annulus to increaseoperational efficiency. In addition, annular momentum of the air streamfrom the compressor is conserved to reduce the overall pressure loss andagain increase in operational efficiency.

Injection of air for film cooling is accomplished through the use ofcooling air openings in one or more of the walls of the annularcombustor.

Where the air film injection is accomplished through the radially innerand/or radially outer walls of the combustor, it is preferablyaccomplished through the provision of a plurality of axially extendingrows of openings while cooling air film injection through the radiallyextending wall of the combustor is accomplished through the use ofradially extending rows of openings.

In either case, elongated cooling strips having a shape somewhat akin tothat of a flattened "S" are utilized. The cooling strips have one edgesecured to the corresponding wall of the annular combustor and theopposite edge spaced therefrom. The opposite edges overlie correspondingones of the rows of cooling air openings and in the case of the radiallyinner and outer walls are axially directed and in the case of theradially extending wall are generally radially directed. The oppositeedges are downstream in the direction of swirl within the annularcombustor from the edges that are attached to the respective walls. As aconsequence, air enter the combustor through the cooling air opening isdirected by the cooling strip in the tangential direction and in closeproximity to the associated wall to thereby generate the cooling airfirm.

According to a preferred embodiment, the cooling air openings are influid communication with the diffuser to receive compressed airtherefrom.

In a highly preferred embodiment, the fuel injectors comprise fuelnozzles having ends within the primary combustion annulus and airatomizing nozzles for the combustion supporting air surround each of theends of the fuel injector fuel nozzles.

The invention contemplates the use of a compressed air housingsurrounding the combustor in spaced relation thereto and in fluidcommunication with the diffuser. The cooling air openings open to theinterface of the housing and combustor to receive compressed airtherefrom.

Other objects and advantages will become apparent from the followingspecification taken in connection with the accompanying drawings.

DESCRIPTION OF THE DRAWINGS

FIG. 1 is a somewhat schematic, fragmentary, sectional view of a turbinemade according to the invention;

FIG. 2 is a fragmentary sectional view taken approximately along theline 2--2 in FIG. 1; and

FIG. 3 is a fragmentary enlarged sectional view of a cooling strip thatmay be used in the invention.

DESCRIPTION OF THE PREFERERD EMBODIMENT

An exemplary embodiment of a gas turbine made according to the inventionis illustrated in the drawings in the form of a radial flow gas turbine.However, the invention is not so limited, having applicability to anyform of turbine or other fuel cornbusting device requiring an annularcombustor.

The turbine includes a rotary shaft 10 journalled by bearings not shown.Adjacent one end of the shaft 10 is an inlet area 12. The shaft 10mounts a rotor, generally designated 14 which may be of conventionalconstruction. Accordingly, the same includes a plurality of compressorblades 16 adjacent the inlet 12. A compressor blade shroud 18 isprovided in adjacency thereto and just radially outwardly of theradially outer extremities of the compressor blades 18 is a conventionaldiffuser 20.

Oppositely of the compressor blades 16, the rotor 14 has a plurality ofturbine blades 22. Just radially outwardly of the turbine blades 22 isan annular nozzle 24 which is adapted to receive hot gasses ofcombustion from a combustor, generally designated 26. The compressorsystem including the blades 16, shroud 18 and diffuser 20 deliverscompressed air to the combustor 26, and via dilution air passages 27 and28, to the nozzle 24 along with the gasses of combustion. That is tosay, hot gasses of combustion from the combustor 26 are directed via thenozzle 24 against the blades 22 to cause rotation of the rotor 14 andthus the shaft 10. The latter may be, of course, coupled to some sort ofapparatus requiring the performance of useful work.

A turbine blade shroud 29 is interfitted with the combustor 26 to closeoff the flow path from the nozzle 24 and confine the expanding gas tothe area of the turbine blades 22.

The combustor 26 has a generally cylindrical inner wall 32 and agenerally cylindrical outer wall 34. The two are concentric and merge toa necked down area 36 which serves as an outlet from the interiorannulus 38 of the combustor to the nozzle 24. A third wall 39, generallyradially extending and concentric with the walls 32 and 34,interconnects the same to further define the annulus 38.

Oppositely of the outlet 36, and adjacent the wall 39, the interiorannulus 38 of the combustor 26 includes a primary combustion zone 40. Byprimary combustion zone, it is meant that this is the area in which theburning of fuel primarily occurs. Other combustion may, in someinstances, occur downstream from the primary combustion area 40 in thedirection of the outlet 36. As mentioned earlier, provision is made forthe injection of dilution air through the passageways 27 and 28 into thecombustor 26 downstream of the primary combustion zone 40 to cool thegasses of combustion to a temperature suitable for application to theturbine blades 22 via the nozzle 24. It should be noted that thepassageways 27 and 28 are configured so that the vast majority ofdilution air flow into the combustor 26 occurs through the passageways28. This, of course, requires the vast majority of dilution air to passabout the generally radially outer wall 34, the third wall 39 and theradially inner wall 32 which in turn provides excellent convectivecooling of these combustor walls and avoids the formation of hot spotson any of the walls 32, 34 and 39.

In any event, it will be seen that the primary combustion zone 40 is anannulus or annular space defined by the generally radially inner wall32, the generally radially outer wall 34 and the wall 39.

A further wall 44 is generally concentric to the walls 32 and 34 and islocated radially outwardly of the latter. The wall 44 extends to theoutlet of the diffuser 20 and thus serves to contain and directcompressed air from the compressor system to the combustor 26.

As best seen in FIG. 2, the combustor 26 is provided with a plurality offuel injection nozzles 50. The fuel injection nozzles 50 have ends 52disposed within the primary combustion zone 40 and which are configuredto be nominally tangential to the inner wall 32. The fuel injectionnozzles 50 generally but not necessary utilize the pressure drop of fuelacross swirl generating orifices (not shown) to accomplish fuelatomization. Tubes 54 surround the nozzles 50. High velocity air fromthe compressor flows through the tubes 54 to enhance fuel atomization.Thus the tubes 54 serve as air injection tubes. When swirl generatingorifices are not used as in the embodiment illustrated, high velocityair flowing through the tubes 34 is the means by which fuel exiting thenozzles 50 is atomized.

The fuel injecting nozzles 50 are equally angularly spaced about theprimary combustion annulus 40 and optionally disposed between each pairof adjacent nozzles 50 there may be a combustion supporting air jet 56.When used, the jets 56 are located in the wall 34 and establish fluidcommunication between the air delivery annulus defined by the walls 34and 44 and the primary combustion annulus 441. These jets 56 may besomewhat colloquially termed "bender" jets as will appear. They are alsooriented so that the combustion supporting air entering through thementers the primary combustion annulus 40 in a direction nominallytangential to the inner wall 32.

Preferably the injectors 50 and jets 56 are coplanar or in relativelyclosely spaced planes remote from the outlet area 36. Such plane orplanes are transverse to the axis of the shaft 10.

When the intended use of the engine requires the delivery of largequantities of bleed air, the wall 44 is provided with a series of outletopenings 58 which in turn are surrounded by a bleed air scroll 60secured to the outer surface of the wall 44. Thus, bleed air to be usedfor conventional purposes may be made available at an outlet (not shown)from the scroll 60.

To prevent the formation of undesirable hot spots on the walls 32, 34and 39 for any of a variety of reasons, the invention contemplates theprovision of means for flowing a cooling air film over the walls 32, 34and 39 on the surfaces thereof facing the annulus 38. Further, theinvention provides means whereby the cooling air film is injected intothe annulus 38 in a generally tangential, as opposed to axial,direction.

Preferably, the injection is provided along each of the walls 32, 34 and39 but in some instances, such injection may occur on less than all ofsuch walls as desired.

In the case of the radially inner wall 32, the same is provided with aseries of apertures 70. Preferably, the apertures 70 are arranged in aseries of equally angularly spaced, generally axially extending rows.Thus, the three apertures 70 shown in FIG. 2 constitute one aperture ineach of three such rows while the apertures 70 illustrated in FIG. 1constitute the apertures in a single such row.

A similar senes of equally angularly spaced, axially extending rows ofapertures 72 is likewise provided in the wall 34.

Similarly, in the case of the wall 39, there are a series of generallyradially extending rows of apertures 74. As can be readily appreciated,the apertures 70, 72 and 74 establish fluid communication between theannulus defined by the wall 44 and the wall 34, a radially extendingannulus defined by the wall 39 and a wall 80 connected to the wall 44,and the connecting annulus defined by the wall 32 and a connecting wall82.

Thus tangential and film-like streams of cooling air enter the annulus38 through the openings 70, 72 and and cooling strips 86, 88, and 90 areapplied respectively to the walls 32, 34 and 39.

As a consequence of this construction, the air flowing in the annuliabout the combustor 26 will remove heat therefrom by external convectivecooling of the walls 32, 34 and 39. Similarly the cooling air film onthe sides of the walls 32, 34 and 39 fronting the annulus 38 resultingfrom film-like air flow into the annulus 38 through the apertures 70, 72and 74 minimizes the input of heat from the flame within the combustor26 to the walls 32, 34 and 39.

Thus, in the preferred embodiment, the entirety of the internal surfaceof all of the walls, 32, 34 and 39 is completely covered with a film ofair. The ability to completely cool the internal walls of a combustor isdifficult to accomplish, particularly as combustor size decreases.However, utilizing the novel technique of tangential injection of air asherein disclosed readily accomplishes the establishment of a completewall covering film to provide improved wall cooling. The film furtherserves to minimize carbon build-up and the elimination of hot spots onthe combustor walls.

These advantages are enhanced by reason of the jets of air which resultfrom air flow through the apertures 70, 72 and 74. Such jets of airimpact upon the cooling strips to cool them. The cooling strips 86, 88and 90 are further cooled by the aforementioned film of air flowing overthem. The cooling strips also act as a local barrier to convective andradiative heating of the walls 32, 34 and 39 by the flame burning withincombustor 26.

The cooling strips 86, 88 and 90 are generally similar one to the otherand accordingly, it is believed that a complete understanding of theoperation of the same can be achieved simply from understanding theoperation of one. Thus, only the cooling strip 86 will be described.

With reference to FIG. 3, the cooling strip 86 is seen to be in theshape of a generally flattened "S" having an upstream edge 92 bonded tothe wall 32 just upstream of a corresponding row of the openings 70 byany suitable means as brazing or, for example, a weld 94. Because of theS shape of the cooling step 86, this results in the opposite ordownstream edge 96 being elevated above the opening 70 with an exitopening 98 being present. The exit opening 98 is elongated in the axialdirection along with the edge 96 and also opens generally tangentiallyto the wall 32. Consequently, air entering the annulus 38 through theopenings in the direction of arrows 100 (FIGS. 2 and 3) will flow in afilm-like fashion in a generally tangential direction along the wall 32on its interior surface to cool the same. The air flow indicated byarrows 102 in FIG. 2 illustrate the corresponding tangential, film-likeflow of cooling air on the interior of the wall 34 while additionalarrows 104 in FIG. 2 illustrated a similar, tangential film-like airflow of air entering the openings 74 in the wall 39.

Operation is generally as follows. Fuel emanating from each of thenozzles 50 will enter along a line such as shown at "F". This line willof course be straight and it will be expected that the fuel will divergefrom it somewhat. Assuming bender jets 56 are used, as the fuelapproaches the adjacent bender jet 56 in the clockwise direction, theincoming air from the diffuser 20 and compressor blades 16 will tend todeflect or bend the fuel stream to a location more centrally of theprimary combustion annulus 40 as indicated by the curved line "S". Therewill, of course, be a substantial generation of turbulence at this timeand such turbulence will promote uniformity of burn within the primarycombustion annulus 40 and this in turn will result in a uniformcircumferential turbine inlet temperature distribution at the nozzle 24and at radially outer ends of the turbine blades 22. Such uniformturbine inlet temperature distribution is achieved in a combustor madeaccording to the invention utilizing many fewer fuel injecting nozzles50 than would be required according to prior art teachings. As a resultof the invention, and even without the use of the bender jets 56,through the use of tangential fuel injection and cooling filmintroduction, a combustor made according to the invention will requireabout half the number of fuel injector nozzles 50 as would aconventional combustor of equal volume. In particular, the two will haveapproximately the same so-called "pattern factor".

If the bender jets 56 are added without adding nozzles 50, animprovement in pattern factor will be obtained over the conventionalcombustor.

In any event, resulting elimination of a number of fuel injector nozzles50 provides a substantial cost savings. Moreover, in engines having anincreased combustor volume, a further substantial reduction in thenumber of fuel injectors by as much as 80% of those required accordingto conventional practice may be obtained.

It will also be observed that where the number of fuel injectionsnozzles 50 is halved using the principals of the invention, the fuelflow passages of the remaining fuel injection nozzles, assuming they arecylindrical, can be increased in diameter slightly over 40%. Thisincrease in diameter reduces the possibility of plugging of the fuelinjection nozzles 50 to provide a more trouble free apparatus. Thischaracteristic of the invention assumes extreme importance in smallengines which utilize small combustors and thus have relatively smallfuel flows, particularly at low engine speeds or while starting at highaltitudes.

In addition, the injection of cooling air in a film-like manner achievedby means of the openings 70, 72 and 74 and associated cooling strips 86,88 and 90 further minimizes the possibility of a hot spot on a wallcoming into existence and thereby prolongs the life of the apparatus.Significantly, the tangential injection of the cooling air film in thesame direction as the swirl within the annulus 38 does not provide anaxial impetus to fuel droplets entering the primary combustion zone 40from the nozzles 50. As a consequence, there is ample time for such fuelto fully and completely vaporize within the primary combustion zone 40and thereby achieve highly efficient combustion. For example, in onecombustor made according to the invention tested at 10% of rated enginespeed with a combustor pressure drop of only 0.8 inches of water, ashort efficient flame was obtained using No. 2 diesel fuel. In contrast,a conventional annular combustor using conventional swirl air blastinject on would typically be unable to sustain combustion under similarcircumstances. Thus, an engine employing the invention is more easilystarted, a feature that may be particularly critical when high altitudeoperation is used as, for example, when the engine is used as part of anauxiliary power unit or an emergency power unit. Because a high degreeof tangential motion or swirl is found desirable in a turbine madeaccordingly to the invention, desire vanes such as those somewhatschematically illustrated at 106 in FIG. 1 may be relatively minimalthereby reducing the complexity of the invention. The swirl that is thuspermitted conserves the angular velocity of the compressed air as itleaves the diffuser 20 so that the pressure drop is minimized, therebyenhancing operational efficiency. Furthermore, since the turbine nozzle24 is desire, need to impart swirl to the hot gases directed against theturbine blades 22, the fact that the gases are already swirling as aresult of tangential air and fuel injection minimizes the directionalchange applied to such gases by the nozzle 24 to provide a furtherincrease in efficiency.

At the same time, the use of minimal deswirl vanes 106 allows theinitial swirl that is typically imparted to the compressed air by thecompressor 16 and diffuser 20 to be retained outside the combustor 26allowing bleed air, which is commonly obtained from a circumferentialvent enclosed by a scroll, to be obtained with a high degree ofefficiency.

According to the invention, the combustor is sized by an equation of theform: ##EQU1## Where K is a constant;

W_(a) is the combustor air flow in pounds per second;

T₃ is the turbine inlet temperature in degrees Rankine;

T₂ is the combustor inlet temperature in degree Rankine;

ΔP/P is the combustor pressure drop×100;

P is the combustor air inlet pressure in psia;

ΔP is the combustor pressure drop in psia;

D is the mean combustor height in inches;

H is the mean combustor width in inches;

N is the number of fuel injectors; and

R is the engine pressure ratio.

The present invention provides a trade-off between combustor volume andthe number of injectors. It is a trade-off that cannot be achieved inconventional combustors. Specifically, in a conventional combustor, thenumber of injectors is determined generally by the expression N=πD/H.

If the number of injectors as defined by the preceding equation isreduced, there is a senous increase in turbine inlet hot spots. In onecombustor made according to the invention, only four injectors wererequired whereas normal practice would require about thirteen suchinjectors. Further, in the combustor made according to the invention, apattern factor of 0.095 was obtained. The pattern factor is a measure ofthe uniformity of temperature throughout the combustion area and isdefined by the formula ##EQU2## where T_(h) is a temperature of thehottest spot in degrees Rankine.

In any event, the pattern factor of 0.095 obtained in a combustor madeaccording to the invention is twice as good as the pattern factor thatwould be obtained in normal practice with thirteen injectors.

Further, when one of the fuel injectors in the four injector structuremade according to the invention was plugged up to simulate a typicalfield failure. the pattern factor increased only to 0.11, a negligibleincrease. Converely, extensive experience in turbine engines hasindicated that if one injector plugs up in a conventional combustor, theresulting hot spot will seriously damage or even destroy the turbineengine.

Similarly, when a combustor employing two diametrically oppositeinjectors with two intermediate bender jets was employed, a patternfactor of 0.2 was obtained. This pattern factor its comparable to thatwhich would be obtained in a conventional combustor utilizing 13injectors. The improvements in pattern factors along with the ability totolerate plugging as well as the elimination of a large number ofinjectors clearly the demonstrates the superiority of the invention.

In addition, in a combustor made according to the invention, a test wasrun with fuel flowing only out of one injector of the four provided. Theinjector from which fuel was flowing was the lowermost one and the testwas to simulate start-up of the engine at very high altitudes when, dueto so-called "manifold head" effects, at low fuel flow rates.Substantially all fuel flows into the combustor through the lowermostinjector. The resulting time visually observed spread about the entirecombustor and the pattern factor was a tolerable 0.33. Conversely, in aconventional combustor wherein fuel is flowed only through one injector,a very localized of lame with inefficient burning is observed andstarting at altitudes is poor.

Thus, in addition to the previously stated advantages, the invention isideally suited for use in turbine engines, particularly small turbineengines, that may be operated at high altitudes and require starting atsuch altitudes as well.

We claim:
 1. A gas turbine comprising:a rotor including compressorblades and turbine blades; an inlet adjacent one side of said compressorblades; a diffuser adjacent the other side of said compressor blades; anozzle adjacent said turbine blades for directing hot gasses at saidturbine blades to cause rotation of said rotor; an annular combustorhaving radially inner and outer walls connected by a generally radiallyextending wall about said rotor and having an outlet connected to saidnozzle and a primary combustion annulus defined by said walls remotefrom said outlet, a plurality of fuel injectors to said primarycombustion annulus and being substantially equally .[.angular.]..Iadd.angularly .Iaddend.spaced therearound and configured to injectfuel into said primary combustion annulus in a nominally tangentialdirection; and means .[.associated with.]. .Iadd.on .Iaddend.each ofsaid walls for injecting a film-like stream of cooling air into saidprimary combustion annulus in a generally tangential direction.
 2. Thegas turbine of claim 1 wherein said .Iadd.cooling air .Iaddend.injectionmeans include cooling air openings in fluid communication with saiddiffuser to receive compressed gas therefrom.
 3. The gas turbine ofclaim 2 wherein a compressed gas housing surrounds said combustor inspaced relation thereto and is in fluid communication with saiddiffuser, said cooling air openings extending to the interface of saidhousing .[.an.]. .Iadd.and said .Iaddend.combustor to receive compressedgas therefrom.
 4. The gas turbine of claim 1 wherein said fuel injectorscomprise fuel nozzles having ends within said primary combustionannulus, and atomizing nozzles for said combustion supporting gassurrounding said ends.
 5. A gas turbine comprisinga rotor includingcompressor blades and turbine blades; an inlet adjacent one side of saidcompressor blades: a diffuser adjacent the other side of said compressorblades: a nozzle adjacent said turbine blades for directing hot gassesat said turbine blades to cause rotation of said rotor, and an annularcombustor about said rotor having an outlet to said nozzle, an innerwall and an outer wall spaced therefrom and a connecting radial walldefining an annular combustion space, a plurality of cooling air filminjectors at substantially equally angularly spaced locations about eachsaid wall and oriented to inject a film-like air stream on theassociated wall generally tangentially to said annular combustion space.6. The gas turbine of claim 5 wherein each said cooling air filminjector comprises a row of openings in the associated wall and acooling strip having an edge overlying and spaced from said row.
 7. Thegas turbine of claim 6 wherein said cooling strips each have a crosssection in the shape of a flatted "S".
 8. The gas turbine of claim 6wherein the rows and strips associated with said inner and outer wallsextend generally axially.
 9. The gas turbine of claim 6 wherein the rowsand strips associated with said radial wall extend generally radially..Iadd.10. A gas turbine comprising:a rotary compressor; a rotary turbinewheel mounted for rotation about an axis and coupled to said compressorto drive the same; a nozzle adjacent said turbine wheel for directinghot gases thereat to rotate the same; an annular combustor about saidturbine wheel having radially inner and outer walls, a radiallyextending wall and an outlet connected to said nozzle and opposite saidradially extending wall with a primary combustion annulus defined bysaid walls remote from said outlet: and a plurality of circumferentiallyspaced fuel injectors adjacent said radially outer wall and having endsdisposed in said primary combustion annulus so as to inject fuelthereinto in a direction nominally tangential to said radially innerwall at radii passing through said axis. .Iaddend. .Iadd.11. The gasturbine of claim 10 wherein said combustor further includes a pluralityof circumferentially spaced air injector tubes mounted on said radiallyouter wall and oriented to direct air into said primary combustionannulus in a direction nominally tangential to said radially inner wall..Iaddend. .Iadd.12. The gas turbine of claim 11 wherein said fuelinjectors are located within at least some of said air injection tubes..Iaddend. .Iadd.13. A gas turbine comprising: a rotary compressor; arotary turbine wheel mounted for rotation about an axis and coupled tosaid compressor to drive the same; a nozzle adjacent said turbine wheelfor directing hot gases thereat to rotate the same; an annular combustorabout said turbine wheel having radially inner and outer walls, aradially extending wall and an outlet connected to said nozzle andopposite said radially extending wall with a primary combustion annulusdefined by said walls remote from said outlet; a plurality ofcircumferentially spaced fuel injectors for injecting fuel into saidprimary combustion annulus in a nominally tangential direction; and aplurality of circumferentially spaced air injector tubes mounted on saidradially outer wall and oriented to direct air into said primarycombustion annulus in a direction nominally tangential to said radiallyinner wall at radii passing through said axis. .Iadd.14. The gas turbineof claim 13 wherein said fuel injectors are located within at least someof said air injection tubes. .Iadd.15. A gas turbine comprising: arotary compressor; a rotary turbine wheel mounted for rotation about anaxis and coupled to said compressor to drive the same; a nozzle adjacentsaid turbine wheel for directing hot gases thereat to rotate the sameabout said axis; an annular combustor about said turbine wheel havingradially inner and outer walls, a radially extending wall and an outletconnected to said nozzle and opposite said radially extending wall witha primary combustion annulus defined by said walls remote from saidoutlet; a plurality of circumferentially spaced fuel injectors adjacentsaid radially outer wall and having ends disposed in said primarycombustion annulus so as to inject fuel thereinto in a directionnominally tangential to said radially inner wall at radii passingthrough said axis; a plurality of circumferentially spaced air injectortubes mounted on said radially outer wall and oriented to direct airinto said primary combustion annulus in a direction nominally tangentialto said radially inner wall at radii passing through said axis; and atleast some of said air injection tubes being in surrounding relation toone of said fuel injector so that air flowing through said air injectiontubes assists in atomizing fuel injected by said injectors. .Iaddend.